The present invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) followed in turn by a low pressure turbine (LPT). The HPT drives the compressor, and the LPT drives an upstream fan in a typical turbofan aircraft engine application.
The HPT first receives the hottest combustion gases from the combustor and must be suitably cooled for ensuring a suitable useful life thereof. The turbine stage includes a turbine stator nozzle disposed at the exit of the combustor for first receiving and guiding the combustion gases therefrom. The nozzle includes a row of hollow stator vanes mounted between outer and inner bands.
The nozzle vanes direct the combustion gases through the first stage of turbine rotor blades extending radially outwardly from a supporting rotor disk. Each turbine blade includes an airfoil over which the combustion gases flow, which are bound by a platform disposed at the root of the airfoil. The platform is joined to a supporting dovetail that mounts the individual turbine blades in corresponding dovetail slots formed in the perimeter of a supporting rotor disk.
Both the stator vanes and rotor blades are hollow and provided with corresponding cooling circuits therein for channeling pressurized air bled from the compressor for cooling these components against the high heat load generated by the hot combustion gases during operation.
The corresponding airfoil configurations of the nozzle vanes and turbine blades are different for the specialized aerodynamic performance thereof. And, the vanes and blades are configured differently and mounted differently for their different operation in the stator nozzle and rotor disk.
Accordingly, the prior art of turbine airfoil cooling for both nozzle vanes and rotor blades is quite crowded, sophisticated, and esoteric for addressing the fundamental differences between vanes and blades and their operating environments in the gas turbine engine for maximizing cooling performance thereof over their varying configurations.
The typical airfoil of both the nozzle vanes and turbine blades includes a generally concave pressure side and an opposite, generally convex suction side extending chordally between opposite leading and trailing edges, and extending radially across the longitudinal span thereof. Each nozzle vane is mounted at its opposite radial ends to corresponding outer and inner bands. Each turbine blade is mounted at its dovetail end to the perimeter of the rotor disk, with the radially outer tip end of the airfoil extending freely in close proximity to the surrounding turbine shroud.
The nozzle vanes and rotor blades accordingly require different cooling circuits therein for the different configurations thereof and the different operating environments to maximize the cooling performance of the limited compressor bleed air channeled therethrough. And, the turbine blades experience the additional complication of rotation during operation which introduces centrifugal forces on the cooling air and secondary Coriolis forces due to the secondary direction of flow turning inside the rotating blades.
Nevertheless, nozzle vanes and turbine blades share similar cooling features such as radially extending flow passages, internal turbulators for heat transfer, film cooling holes arranged in multiple radial rows or columns over the pressure sidewall or suction sidewall, or both, and additional trailing edge outlets for discharging the spent cooling air.
Film cooling is a common cooling practice in which the cooling air is discharged from inside the airfoil in thin films that provide a thermally insulating air blanket over the external surface of the airfoils for protection from the surrounding hot combustion gases.
Inside the corresponding airfoils, impingement cooling techniques may be provided for impingement cooling selected locations of the internal surface of the airfoil against the high heat loads found outside the airfoil. And, the various cooling circuits in the airfoil are typically arranged in independent circuits specifically dedicated for different portions of the airfoil between the leading and trailing edges thereof and along the different pressure and suction sidewalls.
The internal dividing ribs in the airfoils that define the corresponding flow channels are themselves cooled by the cooling air channeled through the flow channels. The pressure and suction sidewalls of the airfoil are directly subjected to the external hot combustion gases which causes them to operate at elevated temperature.
In contrast, the internal ribs are protected from the external combustion gases by the sidewalls themselves and operate at substantially lower temperatures.
Accordingly, the different operating temperatures of the external sidewalls of the airfoils and their internal ribs correspondingly create differential temperatures therebetween that in turn create thermal stress.
Both the nozzle vanes and turbine blades are subject to such differential thermal stress, as well as additional stress from the pressure forces of the combustion gases themselves. And, the turbine blades are additionally subject to centrifugal stress from rotating the blades during operation.
Accordingly, the design of turbine nozzle vanes and rotor blades is remarkably complex with the temperature and stress distribution varying substantially over the opposite pressure and suction sides, both in chord between the leading and trailing edges thereof and radially over the longitudinal span.
And, the temperature and stress also vary in complex distributions inside each turbine airfoil along the corresponding ribs therein that define the various flow channels and flow circuits in each airfoil.
The durability and life of the individual turbine airfoil is therefore limited by the maximum temperature and maximum stress exhibited anywhere over the complex configuration of the airfoil, both outside and inside, which leads to the accumulation of thermal fatigue damage over extended operation in a gas turbine engine.
Accordingly, it is desired to further improve the cooling configuration of a turbine airfoil for further improving its durability and life.